Aircraft hybrid propulsion fan drive gear system dc motors and generators

ABSTRACT

An aircraft propulsion system is disclosed and includes a first gas turbine engine including a first input shaft driving a first gear system, a first fan driven by the first gear system, a first generator supported on the first input shaft and a fan drive electric motor providing a drive input to the first fan, a second gas turbine engine including a second input shaft driving a second gear system, a second fan driven by the second gear system, a second generator supported on the second input shaft and a second fan drive electric motor providing a drive input to the second fan and a controller controlling power output from each of the first and second generators and directing the power output between each of the first and second fan drive electric motors.

CROSS REFERENCE TO RELATED APPLICATION

This application is a divisional of U.S. application Ser. No. 16/039,849filed on Jul. 19, 2018.

BACKGROUND

A gas turbine engine typically includes a fan section, a compressorsection, a combustor section and a turbine section. Air entering thecompressor section is compressed and delivered into the combustionsection where it is mixed with fuel and ignited to generate ahigh-energy exhaust gas flow. The high-energy exhaust gas flow expandsthrough the turbine section to drive the compressor and the fan section.The compressor section typically includes low and high pressurecompressors, and the turbine section includes low and high pressureturbines.

A speed reduction device such as an epicyclical gear assembly may beutilized to drive the fan section such that the fan section may rotateat a speed different than the turbine section so as to increase theoverall propulsive efficiency of the engine. In such enginearchitectures, a shaft driven by one of the turbine sections provides aninput to the epicyclical gear assembly that drives the fan section at areduced speed such that both the turbine section and the fan section canrotate at closer to optimal speeds.

Incorporation of electric power in gas turbine engines is currentlysubstantially limited to accessory components. Advances in electricmotors and generators along with demands for ever increasing engineoperating efficiencies warrant consideration of alternate engineconfigurations.

Geared architectures have improved propulsive efficiency and promptedturbine engine manufacturers to seek further improvements to engineperformance including improvements to propulsive efficiencies.

SUMMARY

An aircraft propulsion system according to an exemplary embodiment ofthis disclosure includes, among other possible things, a first gasturbine engine including a first input shaft driving a first gearsystem, a first fan driven by the first gear system, a first generatorsupported on the first input shaft and a fan drive electric motorproviding a drive input to the first fan, a second gas turbine engineincluding a second input shaft driving a second gear system, a secondfan driven by the second gear system, a second generator supported onthe second input shaft and a second fan drive electric motor providing adrive input to the second fan and a controller controlling power outputfrom each of the first and second generators and directing the poweroutput between each of the first and second fan drive electric motors.

In a further embodiment of the foregoing aircraft propulsion system, thefirst and second gear systems provide a main drive input to acorresponding one of the first and second fans. The first and second fandrive electric motors provide a supplemental drive input to at least oneof the first and second fans.

In a further embodiment of any of the foregoing aircraft propulsionsystem, the controller is configured to balance power generated by boththe first and second generators between the first fan drive electricmotor and the second fan drive electric motor responsive to an imbalancebetween the main drive input of one of the first gas turbine engine andthe second gas turbine engine.

In a further embodiment of any of the foregoing aircraft propulsionsystem, the controller is configured to direct power generated by eachof the first and second generators independent of the other of the firstand second generators to one or both of the first fan drive electricmotor and the second fan drive electric motor.

In a further embodiment of any of the foregoing aircraft propulsionsystem, each of the first generator and the second generator includes arotor supported on the corresponding one of the first and second inputshafts and a stator disposed on a static structure relative to therotor.

In a further embodiment of any of the foregoing aircraft propulsionsystem, each of the first and second generators includes a first statorand a first rotor including a first set of poles providing power to afirst phase of the fan drive electric motor. A second stator and asecond rotor including a second set of poles provides power to a secondphase of the fan drive electric motor.

In a further embodiment of any of the foregoing aircraft propulsionsystem, the first and second fan drive electric motors are electricallycoupled to the corresponding one of the first stator and the secondstator. The first set of poles and the second set of poles are clockedrelative to each other such that rotation of the input shaft commutatesthe first phase and the second phase to drive the fan drive electricmotor.

In a further embodiment of any of the foregoing aircraft propulsionsystem, the first and second fan drive electric motor includes apermanent magnet rotor mounted to a fan shaft. The permanent magnetrotor includes a plurality of poles corresponding with the first set ofpoles and the second set of poles of both the first and secondgenerators.

In a further embodiment of any of the foregoing aircraft propulsionsystem, the first and second gear systems are configured to provide aratio between an input speed of a corresponding one of the first andsecond input shafts and an output speed of the corresponding fan shaft.Commutation of the plurality of poles of the first and second fan driveelectric motors corresponds with the gear ratio and the number of polesin each of the first set of poles and the second set of poles.

In a further embodiment of any of the foregoing aircraft propulsionsystem, for each of the first and second generators, the first rotor isspaced axially apart from the second rotor on the corresponding one ofthe first and second input shafts.

In a further embodiment of any of the foregoing aircraft propulsionsystem, each of the first and second generators includes a third rotorincluding a third set of poles and a third stator. Each of the first andsecond fan drive electric motors includes poles corresponding to each ofthe first, second and third set of poles of each of the first and secondgenerators.

In a further embodiment of any of the foregoing aircraft propulsionsystem, at least one battery coupled to the first and second generatorsis included. The first and second generators provide electric power tocharge at least one battery.

Another aircraft propulsion system according to an exemplary embodimentof this disclosure includes, among other possible things, at least twogas turbine engines each including a fan driven by a fan shaft rotatableabout an engine axis, a fan drive electric motor providing asupplemental drive input to the fan, a gear system driven by an inputshaft and coupled to the fan shaft to provide a main drive input fordriving the fan, and a generator means driven by the input shaftconfigured to generate electric power and for driving the fan driveelectric motor; and a controller controlling power output from thegenerator of each of the at least two gas turbine engines and directingpower to at least one of the fan drive electric motor of the at leasttwo gas turbine engines.

In a further embodiment of the foregoing aircraft propulsion system, thegenerator means includes a first stator and a first rotor including afirst set of poles providing electric power to a first phase of the fandrive electric motor and a second stator and a second rotor including asecond set of poles providing power to a second phase of the fan driveelectric motor.

In a further embodiment of any of the foregoing aircraft propulsionsystem, the first set of poles and the second set of poles are clockedrelative to each other such that rotation of the input shaft commutatesthe first phase and the second phase to drive the fan drive electricmotor.

In a further embodiment of any of the foregoing aircraft propulsionsystem, the generator means includes a third stator and a third rotorincluding a third set of poles providing electric power to a third phaseof the fan drive electric motor. The first, second and third set ofpoles are clocked relative to each other such that rotation of the inputshaft commutates the first phase, second phase and third phase of thefan drive electric motor.

In a further embodiment of any of the foregoing aircraft propulsionsystem, the gear system is configured to provide a ratio between aninput speed of the input shaft and an output speed of the fan shaft.Commutation of the plurality of poles of the fan drive electric motorcorresponds with the gear ratio and the number of poles in each of thefirst set of poles and the second set of poles.

In a further embodiment of any of the foregoing aircraft propulsionsystem, the controller is configured to balance power generated by thegenerator of each of the at least two gas turbine engines to distributeelectric power between to the fan drive electric motor of each of the atleast two gas turbine engines responsive to an imbalance between themain drive input of one of the at least two gas turbine engines.

A method of operating an aircraft propulsion system according to anexemplary embodiment of this disclosure includes, among other possiblethings, generating electric energy with a generator mounted to an inputshaft driving a gear system at a first speed in each of a first gasturbine engine and a second gas turbine engine, driving a fan shaft ofeach of the first and second gas turbine engine at a second speeddifferent than the first speed with a primary rotational input from thecorresponding gear system and driving the fan shaft with a supplementalrotational input with a fan drive electric motor driven by electricenergy generated by the generator disposed on the input shaft of atleast one of the first and second gas turbine engines and distributingelectric power generated from the generator of each of the first andsecond gas turbine engines through an electric network to at least onefan drive electric motor of the first and second turbine engines.

In a further embodiment of the foregoing method of operating an aircraftpropulsion system, the generator comprises a first generator portionproviding electric power to a first phase of the fan drive electricmotor and a first battery, and a second generator providing electricpower to a second phase of the fan drive electric motor and a secondbattery.

In a further embodiment of any of the foregoing methods of operating anaircraft propulsion system, the generator includes a third generatorportion providing electric power to a third phase of the fan driveelectric motor and third battery.

In a further embodiment of any of the foregoing methods of operating anaircraft propulsion system, commutating electric power provided to atleast the first phase and second phase of the electric motor by clockinga first set of poles of first generator portion relative to a second setof poles of the second generator portion is included.

In a further embodiment of any of the foregoing methods of operating anaircraft propulsion system, engaging the supplemental rotational inputwith the fan drive electric motor responsive to a decreased load on afan drive turbine is included to drive the input shaft of one of thefirst and second gas turbine engines with electric power from agenerator of each of the first and second gas turbine engines.

In a further embodiment of any of the foregoing methods of operating anaircraft propulsion system, distributing electric power to the fan driveelectric motor of each of the first and second gas turbine enginesresponsive to different loads on the fan drive turbine to balance loadson the fan drive turbine between the first and second gas turbineengines.

In a further embodiment of any of the foregoing methods of operating anaircraft propulsion system, commutating the electric power provided toat least the first phase and second phase of the electric motor isincluded by charging and discharging at least one of a first battery anda second battery.

Although the different examples have the specific components shown inthe illustrations, embodiments of this invention are not limited tothose particular combinations. It is possible to use some of thecomponents or features from one of the examples in combination withfeatures or components from another one of the examples.

These and other features disclosed herein can be best understood fromthe following specification and drawings, the following of which is abrief description.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a schematic view of an example gas turbine engine.

FIG. 2 is a schematic view of an example aircraft propulsion system.

FIG. 3 is a schematic view of a fan drive system of one engine of theaircraft propulsion system.

FIG. 4 is a schematic view of components of the example fan drivesystem.

FIG. 5 is a schematic view of components of another example fan drivesystem.

FIG. 6 is a schematic view of components of yet another example fandrive system.

FIG. 7 is a schematic view of another example fan drive system.

FIG. 8 is a schematic view of one example operating mode of the exampleaircraft propulsion system.

FIG. 9 is a schematic view another example operating mode of the exampleaircraft propulsion system.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. The fan section 22 drivesair along a bypass flow path B in a bypass duct defined within a nacelle18, and also drives air along a core flow path C for compression andcommunication into the combustor section 26 then expansion through theturbine section 28. Although depicted as a two-spool turbofan gasturbine engine in the disclosed non-limiting embodiment, it should beunderstood that the concepts described herein are not limited to usewith two-spool turbofans as the teachings may be applied to other typesof turbine engines including three-spool architectures.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects, a first (or low) pressure compressor 44 and a first (orlow) pressure turbine 46. The inner shaft 40 is connected to a fansection 22 through a speed change mechanism, which in exemplary gasturbine engine 20 is illustrated as a geared architecture 48 to drivefan blades 42 at a lower speed than the low speed spool 30. The highspeed spool 32 includes an outer shaft 50 that interconnects a second(or high) pressure compressor 52 and a second (or high) pressure turbine54. A combustor 56 is arranged in exemplary gas turbine 20 between thehigh pressure compressor 52 and the high pressure turbine 54. Amid-turbine frame 58 of the engine static structure 36 may be arrangedgenerally between the high pressure turbine 54 and the low pressureturbine 46. The mid-turbine frame 58 further supports bearing systems 38in the turbine section 28. The inner shaft 40 and the outer shaft 50 areconcentric and rotate via bearing systems 38 about the engine centrallongitudinal axis A which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 58 includes airfoils 60 whichare in the core airflow path C. The turbines 46, 54 rotationally drivethe respective low speed spool 30 and high speed spool 32 in response tothe expansion. It will be appreciated that each of the positions of thefan section 22, compressor section 24, combustor section 26, turbinesection 28, and fan drive gear system 48 may be varied. For example,gear system 48 may be located aft of the low pressure compressor 44 andthe fan blades 42 may be positioned forward or aft of the location ofthe geared architecture 48 or even aft of turbine section 28.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), thegeared architecture 48 is an epicyclic gear train, such as a planetarygear system or other gear system, with a gear reduction ratio of greaterthan about 2.3 and the low pressure turbine 46 has a pressure ratio thatis greater than about five. In one disclosed embodiment, the engine 20bypass ratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout five 5:1. Low pressure turbine 46 pressure ratio is pressuremeasured prior to inlet of low pressure turbine 46 as related to thepressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle. The geared architecture 48 may be an epicycle geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3:1 and less than about 5:1. Itshould be understood, however, that the above parameters are onlyexemplary of one embodiment of a geared architecture engine.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet (10,668 meters). The flight condition of 0.8 Mach and35,000 ft (10,668 meters), with the engine at its best fuelconsumption—also known as “bucket cruise Thrust Specific FuelConsumption (‘TSFC’)”—is the industry standard parameter of lbm of fuelbeing burned divided by lbf of thrust the engine produces at thatminimum point. “Low fan pressure ratio” is the pressure ratio across thefan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The lowfan pressure ratio as disclosed herein according to one non-limitingembodiment is less than about 1.45. “Low corrected fan tip speed” is theactual fan tip speed in ft/sec divided by an industry standardtemperature correction of [(Tram °R)/(518.7 °R)]^(0.5). The “Lowcorrected fan tip speed” as disclosed herein according to onenon-limiting embodiment is less than about 1150 ft/second (350.5meters/second).

The example gas turbine engine includes the fan section 22 thatcomprises in one non-limiting embodiment less than about 26 fan blades42. In another non-limiting embodiment, the fan section 22 includes lessthan about 20 fan blades 42. Moreover, in one disclosed embodiment thelow pressure turbine 46 includes no more than about 6 turbine rotorsschematically indicated at 34. In another non-limiting exampleembodiment, the low pressure turbine 46 includes about 3 turbine rotors.A ratio between the number of fan blades 42 and the number of lowpressure turbine rotors is between about 3.3 and about 8.6. The examplelow pressure turbine 46 provides the driving power to rotate the fansection 22 and therefore the relationship between the number of turbinerotors 34 in the low pressure turbine 46 and the number of blades 42 inthe fan section 22 disclose an example gas turbine engine 20 withincreased power transfer efficiency.

Changes in environmental conditions can require constant adaptations andadjustments to engine operation to provide a desired propulsive output.For example, fuel flow to the combustor 56 may be adjusted dependingboth on a desired propulsive power output and input airflowcharacteristics including pressure and temperatures. Changes in inputairflows may change during operation and require adjustment of fuel flowto maintain the desired propulsive output. There is a certain lagbetween the adjustment and obtaining the operating propulsive output.Although very brief, the lag can affect engine efficiency.

Moreover, changes in power provided by the low pressure turbine 46driving fan section 22 also will add power to the low pressurecompressor 44 and thereby complicate operation. The low pressurecompressor 44 matches operation to that of the high pressure compressor52 and thereby any adjustment to one results in changes to the other.Excessive power input into the low pressure compressor 44 may requirethat air flow be bled off in order to properly match operation of thehigh pressure compressor 52.

The disclosed gas turbine engine 20 includes a fan drive system 62 thatincludes a fan drive electric motor 66 that is driven by a generator 64driven by the low shaft 40. The generator 64 generates electric powerthat is controlled by a controller 68 to add power to the fan section 22parallel to the power provided through the geared architecture 48.Accordingly, the fan drive system 62 enables additional power to beadded to drive the fan section 22 to supplement power provided throughmechanical means from the fan drive low pressure turbine 46. The fandrive electric motor 66 is more responsive to changes in power demandand thereby can reduce fluctuation in fan operation generated by lags inpower transmitted through the mechanical means of the gearedarchitecture 48.

Referring to FIG. 2 with continued reference to FIG. 1, an aircraftpropulsion system 250 for an aircraft 210 is schematically shown andincludes a first gas turbine engine 20A and a second gas turbine engine20B. Each of the first and second gas turbine engines 20A-B areconfigured substantially the same and include corresponding electric fandrive systems 62A-B that are controlled by the common controller 68. Thefirst engine 20A includes a first generator 64A that is driven by afirst gear system 48A to provide power to a first fan drive electricmotor 66A. The second engine 20B includes a second generator 64B that isdriven by a second gear system 48B to provide power to a first fan driveelectric motor 66A.

The first and second generators 64A-B along with the first and secondfan drive electric motors 62A-B are electrically coupled through anelectric network 132 such that electric power may be distributed fromboth of the generators 64A-B to each of the fan drive electric motors62A-B individually, separately and/or proportionally as needed toprovide power to rotate the corresponding fan section 22. It should beappreciated, that although two gas turbine engines are shown anddescribed by way of example, that other multiples of gas turbine engines20A-B could be utilized including three engines, four engines or more.

The example controller 68 includes an aircraft controller 116 thatreceives information from a first full authority digital enginecontroller (FADEC) 112A of the first engine 20A and a second FADEC 112Bof the second engine 20B. The controller 68 utilizes operationalinformation obtained from each of the first and second engines 20A-B tobalance electric power provided to each of the fan drive electric motors66A-B. Power is provided through the electric network 132 that includeselectrical power communication buses 120A-B for each generator 64A-B andelectric power communication buses 118A-B for each fan drive electricmotor 66A-B. The example electric power network 132 is shown andillustrated in this disclosed example schematically and include theelectrical and control connections required to transfer power from thegenerators 64A-B to the fan drive electric motors 66A-B as is understoodby those understanding electric motor and generator operation.

Referring to FIG. 3 with continued references to FIG. 2, the first gasturbine engine 20A is shown schematically separate from the entirepropulsion system 250 to illustrate and explain operation of theelectric fan drive system 62A. Each fan drive system 62A-B for eachengine 20A-B operates in this disclosed example in the same manner.

The fan drive system 62A receives a primary or first drive input 25 fromthe fan drive turbine 46. A supplemental or second drive input 35 isprovided by the fan drive electric motor 66A. The supplemental driveinput 35 provides additional power to the fan section 22 to, among otherpossible things, accommodate variations in power output through thefirst drive input 25. The example fan drive system 62A exploits thespeed relationship between the input shaft 72 and the fan shaft 70inherent in the gear system 48A. The relative speeds and pole counts ofthe electric motor 66A and the generator 64A are made complimentary toprovide a parallel electric power path to drive the fan section 22.

The relationships between the speeds of the input shaft 72 relative tothe fan shaft 70 along with the pole counts are common between each ofthe engines 20A-B of the propulsion system 250 shown in FIG. 2. In otherwords, the number of rotors, pole counts of the first generator 64A notonly corresponds with the first motor 66A of the first engine 20A, butalso the second motor 66B of the second engine 20B. Similarly, thesecond generator 64B is configured to include a number of rotors andpole counts that correspond with both the first motor 66A and the secondmotor 66B. The common configuration between motors 66A-B and thegenerators 64A-B is provided by utilizing the same gear systems 48A andenables cross operation between engines 20A-B.

The example fan drive electric motor 66A is mounted directly to a fanshaft 70 driven by the geared architecture 48A. The geared architecture48A includes a sun gear 74 driven by the input shaft 72 driven by thelow pressure turbine 46. The sun gear 74 drives intermediate gears 76supported by a carrier 80. The intermediate gears 76 rotate within aring gear 78 that is fixed to the engine static structure 36. Thecarrier 80 is coupled to drive the fan shaft 70. The disclosed gearedarchitecture 48 may be referred to as a planetary gear system andprovides a speed reduction ratio between the input shaft 72 and the fanshaft 70 that is equal to 1+ the gear ratio. In this example, the gearedarchitecture 48 has a gear ratio of three (3.0) and therefore the speedreduction is 1+3.0=4.0.

Although a specific mounting configuration is disclosed by way ofexample, the fan drive electric motor 66 may be mounted in an alternateconfiguration that maintains the relative speed relationship the fanshaft 70 and the input shaft 72 and remain within the contemplation andscope of this disclosure.

The example fan drive electric motor 66 includes a permanent magnetrotor 94 mounted to drive the fan shaft 70. A first stator phase 98 anda second stator phase 96 are fixed relative to the rotor 94 on a portionof the engine static structure 36. The electric motor 66 is drivendirectly by electric power produced by the generator 64. The number andconfiguration of phases are disclosed by way of example and othernumbers of phases as understood for electric motor and generatoroperation are within the contemplation and scope of this disclosure.

The example generator 64A includes a first generator portion 92 and asecond generator portion 90. In this disclosed example, the firstgenerator portion 92 is spaced axially apart from the second generatorportion 90. Each of the generator portions 92, 90 provide power to acorresponding one of the first and second stator phases 96, 98 of thefan drive electric motor 66. A controller 68 controls the communicationof electric power from the generator portion 92, 90 to the electricmotor 66.

Referring to FIG. 4, with continued reference to FIG. 3, the firstgenerator portion 92 includes a first rotor 84 mounted to an input shaft72. The input shaft 72 is driven by the fan drive turbine that in thisexample is the low pressure turbine 46. The second generator portion 90includes a second rotor 88 that is also mounted to the input shaft 72.

The first rotor 84 includes a plurality of poles 104. Each of the poles104 is a permanent magnet. In this example, four poles 104 are provided.The second rotor 88 includes a second plurality of poles 106. The secondplurality of poles 106 includes four poles that are clocked relative tothe first plurality of poles 104. In this disclosure the term clocked isutilized to describe the circumferential offset between the first set ofpoles 104 and the second set of poles 106. In this example, the secondset of poles 106 is clocked 90 degrees relative to the first set ofpoles 104.

The relative radial clocking along with the number of poles in each ofthe first and second stators 86, 82 is combined with the relative speedsbetween the input shaft 72 and the fan shaft 70 to provide the requiredcommutation between the first and second stator phases 98, 96. In thisdisclosed example and as discussed above, the speed reduction ratiobetween the speed of the input shaft 72 and the speed of the fan shaft70 is four (4.0). The number of poles in each of the first and secondstators 86, 82 is therefore four or a multiple of four to provide thecorresponding commutation. Moreover, each of the stators 86, 82 willinclude more poles than that provided on the corresponding rotor 88, 84.

The rotor 94 of the fan drive electric motor 66 includes a plurality ofpoles 95. Each of the poles 95 is a permanent magnet. In this example,sixteen (16) poles are indicated. The number of permanent magnet poles95 is four times (4×) the number of poles 104 of the first rotor 84 andfour times the number of poles 104 of the second rotor 88 as the speedreduction ratio between the input shaft 72 and the speed of the fanshaft 70 is four (4.0). The number of poles of the first stator phase 98is the same as the number of poles of first stator 86. The number ofpoles of the second stator phase 96 is the same as the number of polesof the second stator 82. Although depicted as two phases in thedisclosed non limiting embodiment it should be understood that theconcepts described herein are not limited to two phases as the teachingmay be applied to other electric machine phase configurations includingthree or more phases in the generator 64 and motor 66 with pole countscommensurate with the speed reduction ratio of the geared architecture48. As appreciated, for different gear ratios, different numbers ofpoles would be utilized to provide the required commutation to drive theelectric motor 66.

Each generator portion 92, 90 provides the required phase shift neededto turn the rotor 94 of the fan drive electric motor 66. The relativeclocking between the generator rotors 88, 84 and the motor rotor 94combined with the relative speeds of the fan shaft 70 and the inputshaft 72 provide a mechanical commutation. Accordingly, an electroniccontroller or commutator is not necessary to control operation of themotor 66. The controller 68 is provided to selectively turn the electricmotor 66 on and off and to adjust an amount of power supplied, but isnot utilized in this example embodiment as a commutator.

In operation, the generator 64 rotates with the input shaft 72. Theinput shaft 72 drives the geared architecture 48 to provide a first orprimary rotational input to the fan shaft 70 and thereby the fan section22. The fan shaft 70 is rotated at a second speed that is different and,in this example, less than a first speed of the input shaft 72. Thegenerator 64 provides a first phase of electric power schematicallyindicated at 100 to the first stator phase 98 of the electric motor 66.A second phase of electric power indicated at 102 is provided to asecond phase 96 of the electric motor 66. Commutation between thegenerator stator phases 86, 82 is provided by relative clocking betweenthe generator phases combined with the relative speed between the inputshaft 72 and the fan shaft 70. The first stator phase 86 of thegenerator 64 provides power to the first stator phase 98 of the electricmotor 98. The second stator phase 82 provides power to the second statorphase 96 ninety degrees apart from the first stator phase 86 to providethe required commutation to drive the rotor 94. Although depicted as twophases in the disclosed non limiting embodiment it should be understoodthat the concepts described herein are not limited to two phases as theteaching may be applied to other electric machine phase configurationsincluding three or more phases in the generator 64 and motor 66.

The supplemental input 35 provided by the electric motor 66 isimplemented responsive to variable load demand on the fan drive turbine46. The supplemental power input 35 provided by the motor 66 increasesthe load on the fan drive turbine 46 necessary to meet the decreasedload needed to maintain the desired fan speed and propulsive output.Additionally, the supplemental input 35 can be disengaged to reducepower on the fan drive turbine 46 when a demanded load is increased.

The controller 68 governs operation of the fan drive system 62 andengagement of the supplemental input 35. The controller 68 can be aseparate controller 68 or part of the overall engine and/or aircraftcontroller.

Referring to FIG. 5 with continued reference to FIGS. 3 and 4, first andsecond batteries 110 and 108 are electrically coupled to thecorresponding first and second generator portions 92, 90. The batteries110 and 108 can supplement generated power and also enable storage ofelectric power not used to drive the electric motor 66. The supplementalpower input 35 provided to the electric motor 66 by battery 110 andbattery 108 is implemented responsive to variable load demand on the fandrive turbine 46. The supplemental power input 35 provided by thebattery 110 and battery 108 decreases the load on the fan drive turbine46 necessary to meet the increased load needed to maintain the desiredfan speed and propulsive output. Additionally, the supplemental input 35can be engaged to reduce power on the fan drive turbine 46 when ademanded load is increased.

Referring to FIG. 6, a third generator portion 130 is shown to provide athird phase to drive the fan drive electric motor 66. The thirdgenerator portion 130 includes a third rotor 122 with a third set ofpoles 122. A third stator 124 is provided to communicate to a thirdphase 128 of the electric motor 66. The third rotor 122 is spacedaxially apart along the input shaft 72 next to the first and secondrotors 88, 84. The first, second and third rotors 88, 84 and 130 areclocked relative to each to provide the necessary commutation to drivethe rotor 94 of the electric motor 66. As appreciated, a three phasegenerator and/or electric motor includes phases that are separated bysixty degrees. As with the previous example generator and electric motorexamples, the number of poles in each of the first, second and thirdrotors 88, 84 and 130 are determined relative to the different speedsprovided by the geared architecture 48 to provide the requiredcommutation mechanically by the relative physical structures instead ofusing an electrical commutation controller. The different speeds of theinput shaft and output shaft provided by the geared architecture areutilized to simplify the generation and distribution of power.

Referring to FIG. 7, with continued reference to FIGS. 4 and 5, anotherembodiment of the fan drive system 62 includes another gearedarchitecture 45 that is configured such that the fan shaft 70 is coupledto the ring gear 78. The example geared architecture 45 is sometimesreferred to as a “star” gear system. The geared architecture 45 providesa speed reduction ratio between the shaft 72 and the fan shaft 70 thatis equal to the gear ratio. In this example, the geared architecture 45has a gear ratio of three (3.0) and therefore the speed ratio is 3.0.The number of permanent magnet poles 95 is three times (3×) the numberof poles of the first rotor 84 and three times the number of poles ofthe second rotor 88 as the speed reduction ratio between the input shaft72 and the speed of the fan shaft 70 is three (3.0). The number of polesof the first stator phase 98 is the same as the number of poles of firststator 86. The number of poles of the second stator phase 96 is the sameas the number of poles of the second stator 82. The number of poles ineach of the first and second stators 86, 82 is therefore three or amultiple of three to provide the corresponding commutation.

In the geared architecture 45, the fan shaft 70 is driven in an oppositedirection compared to the input shaft 72 and therefore a phase shift ofpower provided by the first and second generators 92, 90 is required toprovide proper rotation of the motor 66. The voltage polarity acrossfirst stator phase 98 of the electric motor 66 is connectively inverted.The voltage polarity of the second phase of electric power indicated at102 is connected inverted in polarity to the second phase 96 of theelectric motor 66.

Implementation of supplemental power input 35 opposes the rotation offan blades 42. The voltage polarity across first stator phase 98 of theelectric motor 66 is as shown in FIG. 3. The voltage polarity of thesecond phase of electric power indicated at 102 is connected withindicated polarity to the second phase 96 of the electric motor 66 asshown in FIG. 3. The supplemental input 35 provided by the electricmotor 66 is implemented responsive to variable load demand on the fandrive turbine 46. Disengaging the supplemental input 35 removes thetorque provided by electric motor 66 that opposes rotation of fan blades42 and increases a reduction of power on the fan drive turbine 46 when ademanded load is increased.

Referring to FIG. 8 with continued reference to FIG. 2, the disclosedaircraft propulsion system 250 is operable to balance power loadsbetween each of the engines 20A-B to reduce the load on the fan driveturbine 46 and to accommodate operating imbalances between engines. Theelectrical network 132 enables communication of power between any of thefirst and second generators 64A-B and each of the motors 66A. Thecommunication pathway provided by the electric network 132 enablesoperation of the fan drive electric motors 66A-B with power from eitheror both of the generators 64A-B completely or in a proportional manner.

During operation the fan drive turbine 46A-B of each of the engines20A-B provides a first primary drive input 25 is provided by the fandrive turbine 46A-B of each of the engines 20A-B. The primary driveinput 25 drives the geared architecture 48A-B that in turn rotates thatfan 22A-B at speed determined by the gear ratio. The fan drive turbine46A-B also drive the generators 64A-B that produce electric power todrive the corresponding fan drive electric motors 66A-B. The fan driveelectric motors 66A-B provide a supplemental power input schematicallyindicated by arrows 35A-B.

In one example method of operating the disclosed aircraft propulsionsystem 250, both engines 20A-B are operating substantially the same. Inthis example, substantially the same means that both fan drive turbines46A-B are providing a primary drive input 25 for a similar input that issubstantially the same. The input for the fan drive turbines 46A-Bcorrespond with fuel being burned. In the example shown in FIG. 8, bothfan drive turbines 46A-B are generating the same power output given thesame fuel input. As appreciated, the same operation is within expectedtolerances and is not identical. Each of the generators 64A-B provide asupplemental input 35A-B to the corresponding fan drive electric motors66A-B in order to provide the total thrust produced by the engine 20A-B.The total thrust is the propulsive thrust generated by the engine 20A-Bas is understood by those skilled in the art. In this example, thethrust is produced by rotation of the fan 22A-B without consideration ofthe thrust produced by exhaust from the fan drive turbine 46A-B. In thisexample disclosed in FIG. 8, the first primary drive input 25A-B and thesecond supplemental input 35A-B are substantially the same within anexpected range to produce a common amount of thrust.

Referring to FIG. 9, the engines 20A-B are shown schematically anddiffer in efficiencies such that the fan drive turbine 46B does notgenerate the same power as compared to the fan drive turbine 46A of thefirst engine 20A for the same given input. Accordingly, without a meansto supplement power provided to the fan 22B, additional input would beneeded to provide a balance thrust between both engines 20A-B. Theadditional input is added fuel flow through the combustor to drive theturbine 46B at the increased power levels needed to provide balancebetween the engines 20A-B. However, because the generators 64A-B andmotors 66A-B are coupled electrically, power from the first engine 20Acan be transferred to the second engine 20B to provide an increasedamount of supplemental power 35 to power the fan 22B without providingadditional input to the fan drive turbine 46B. In this example, thesupplemental input 35A from the first generator 64A is combined with thesupplemental input 35B from the second generator 64B to increase theadditional power to the electric motor 66B. The additional supplementalpower provided by both supplemental inputs 35A-B can enable balancedthrust production without increasing fuel flow to the fan drive turbine46B.

It should be appreciated, that although the disclosed example isdescribed as providing all power from both generators 64A-B, any partialportion of power from both generators 64A-B could be utilized andcontrolled by the controller 68 to balance loads and provide the desiredthrust balance between the engines 20A-B.

Accordingly, the example aircraft propulsion system 250 includes theelectric network 132 that enables power to be controlled and directedbetween generators of different engines as needed to balance operationacross multiple engines. The parallel inputs of power through the gearedarchitecture 48 and the electric motor 66 enables the application ofpower to the fan section 22 to smooth variations in low pressurecompressor power fluctuations and lag while also reducing dependence oncompressor bleed to match compressor operation.

Although an example embodiment has been disclosed, a worker of ordinaryskill in this art would recognize that certain modifications would comewithin the scope of this disclosure. For that reason, the followingclaims should be studied to determine the scope and content of thisdisclosure.

What is claimed is:
 1. An aircraft propulsion system comprising: a firstgas turbine engine including a first input shaft driving a first gearsystem, a first fan driven by the first gear system, a first generatorsupported on the first input shaft and a fan drive electric motorproviding a drive input to the first fan, wherein the first fan driveelectric motor is electrically coupled and clocked relative to the firstgenerator such that rotation of the first input shaft provides forcommutation of phases of the first electric motor; a second gas turbineengine including a second input shaft driving a second gear system, asecond fan driven by the second gear system, a second generatorsupported on the second input shaft and a second fan drive electricmotor providing a drive input to the second fan, wherein the second fandrive electric motor is electrically coupled and clocked relative to thesecond generator such that rotation of the second input shaft providesfor commutation of phases of the second fan drive electric motor; and acontroller controlling power output from each of the first and secondgenerators and directing the power output between each of the first fandrive electric motor and the second fan drive electric motor.
 2. Theaircraft propulsion system as recited in claim 1, wherein the first gearsystem and the second gear system provide a main drive input to acorresponding one of the first fan and the second fan and the first fandrive electric motor and the second fan drive electric motor provide asupplemental drive input to at least one of the first fan and the secondfan.
 3. The aircraft propulsion system as recited in claim 1, whereinthe controller is configured to balance power generated by both thefirst generator and the second generator between the first fan driveelectric motor and the second fan drive electric motor responsive to animbalance between the main drive input of one of the first gas turbineengine and the second gas turbine engine.
 4. The aircraft propulsionsystem as recited in claim 3, wherein the controller is configured todirect power generated by each of the first generator and the secondgenerator independent of the other of the first generator and the secondgenerator to one or both of the first fan drive electric motor and thesecond fan drive electric motor.
 5. The propulsion system as recited inclaim 1, wherein each of the first generator and the second generatorincludes a rotor supported on the corresponding one of the first inputshaft and second input shaft and a stator disposed on a static structurerelative to the rotor.
 6. The aircraft propulsion system as recited inclaim 5, wherein each of the first generator and second generatorsincludes a first stator and a first rotor including a first set of polesproviding power to a first phase of the fan drive electric motor and asecond stator and a second rotor including a second set of polesproviding power to a second phase of the fan drive electric motor. 7.The aircraft propulsion system as recited in claim 6, wherein each ofthe first fan drive electric motor and the second fan drive electricmotor are electrically coupled to the corresponding one of the firststator and the second stator and the first set of poles and the secondset of poles are clocked relative to each other such that rotation ofthe corresponding one of the first input shaft and the second inputshaft provides for commutation between the first phase and the secondphase to drive the corresponding one of the first fan drive electricmotor and the second fan drive electric motor.
 8. The aircraftpropulsion system as recited in claim 7, wherein each of the first fandrive electric motor and the second fan drive electric motor include apermanent magnet rotor mounted to a corresponding one of a first fanshaft and a second fan shaft, the permanent magnet rotor including aplurality of poles corresponding with the first set of poles and thesecond set of poles of both the first generator and the secondgenerator.
 9. The aircraft propulsion system as recited in claim 8,wherein each of the first gear system and the second gear system areconfigured to provide a ratio between an input speed of a correspondingone of the first input shaft and the second input shaft and an outputspeed of the corresponding first fan shaft and the second fan shaft andthe commutation of the plurality of poles of the first fan drive motorand the second fan drive electric motor corresponds with the gear ratioand the number of poles in each of the first set of poles and the secondset of poles.
 10. The aircraft propulsion system as recited in claim 6,wherein for each of the first generator and the second generator, thefirst rotor is spaced axially apart from the second rotor on thecorresponding one of the first and second input shafts.
 11. The aircraftpropulsion system as recited in claim 9, wherein each of the firstgenerator and the second generator includes a third rotor including athird set of poles and a third stator and each of the first fan driveelectric motor and the second fan drive electric motor includes polescorresponding to each of the first, second and third set of poles ofeach of the first and second generators.
 12. The aircraft propulsionsystem as recited in claim 1, at least one battery electrically coupledto and charged by at least one of the first generator and the secondgenerator.
 13. An aircraft propulsion system comprising at least two gasturbine engines each including, a fan driven by a fan shaft rotatableabout an engine axis, a fan drive electric motor providing asupplemental drive input to the fan, a gear system driven by an inputshaft and coupled to the fan shaft to provide a main drive input fordriving the fan, and a generator means driven by the input shaftconfigured to generate electric power and for driving the fan driveelectric motor, wherein each fan drive electric motor is electricallycoupled and clocked relative to the generator means such that rotationof the input shaft commutates phases of the fan drive electric motor; atleast one battery electrically coupled to the generator means, whereinthe generator means provides electric power to charge the at least onebattery; and a controller controlling power output from the generator ofeach of the at least two gas turbine engines and directing power to atleast one fan drive electric motor of the at least two gas turbineengines.
 14. The aircraft propulsion system as recited in claim 13,wherein the generator means includes a first stator and a first rotorincluding a first set of poles providing electric power to a first phaseof the fan drive electric motor and a second stator and a second rotorincluding a second set of poles providing power to a second phase of thefan drive electric motor.
 15. The aircraft propulsion system as recitedin claim 14, wherein the first set of poles and the second set of polesare clocked relative to each other such that rotation of the input shaftcommutates the first phase and the second phase to drive the fan driveelectric motor.
 16. The aircraft propulsion system as recited in claim15, wherein the generator means includes a third stator and a thirdrotor including a third set of poles providing electric power to a thirdphase of the fan drive electric motor and the first set of poles, thesecond set of poles and the third set of poles are clocked relative toeach other such that rotation of the input shaft commutates the firstphase, second phase and third phase of the fan drive electric motor. 17.The aircraft propulsion system as recited in claim 15, wherein the gearsystem is configured to provide a ratio between an input speed of theinput shaft and an output speed of the fan shaft and commutation of thefirst set of poles and the second set of poles of the fan drive electricmotor corresponds with the gear ratio and the number of poles in each ofthe first set of poles and the second set of poles.
 18. The aircraftpropulsion system as recited in claim 13, wherein the controller isconfigured to balance power generated by the generator means of each ofthe at least two gas turbine engines to distribute electric powerbetween to the fan drive electric motor of each of the at least two gasturbine engines responsive to an imbalance between the main drive inputof one of the at least two gas turbine engines.